This invention relates generally to gas turbine engines and more particularly to turbine nozzle vane airfoils used in such engines.
A gas turbine engine includes a compressor that provides pressurized air to a combustor wherein the air is mixed with fuel and ignited for generating hot combustion gases. These gases flow downstream to one or more turbines that extract energy therefrom to power the compressor and provide useful work such as powering an aircraft in flight. In a turbofan engine, which typically includes a fan placed at the front of the core engine, a high pressure turbine powers the compressor of the core engine. A low pressure turbine is disposed downstream from the high pressure turbine for powering the fan. Each turbine stage commonly includes a stationary turbine nozzle followed in turn by a turbine rotor. The turbine rotor comprises a row of rotor blades mounted to the perimeter of a rotor disk that rotates about the centerline axis of the engine. The nozzle, which channels combustion gases into the turbine rotor in such a manner that the turbine rotor can do work, includes a plurality of circumferentially spaced apart vanes radially aligned with the rotor blades. Turbine nozzles are typically segmented around the circumference thereof to accommodate thermal expansion. Each nozzle segment has one or more nozzle vanes disposed between inner and outer bands that define the radial flowpath boundaries for the hot combustion gases flowing through the nozzle.
The high pressure turbine nozzle is mounted at the exit of the combustor and is therefore exposed to extremely high temperature combustion gases. Thus, the turbine blades and nozzle vanes typically employ internal cooling to keep their temperatures within certain design limits. The nozzle vanes are hollow airfoils having a pressure side wall and a suction side wall joined together at leading and trailing edges. Various conventional configurations exist for cooling both the vanes and the bands. The most common types of cooling include impingement and film cooling. To effect impingement cooling, the airfoil includes one or more perforated hollow inserts that are suitably mounted therein. Cooling air (ordinarily bled off from the engine""s compressor) is channeled into the inserts and then impinges against the inner surface of the airfoil for impingement cooling thereof. Film cooling is accomplished by passing the cooling air through small film cooling holes formed in the airfoil so as to produce a thin layer of cooling air on the outer surface of the airfoil.
During operation, the hot combustion gases flow around each of the nozzle vanes between the outer and inner bands. Accordingly, the turbine nozzle thermally expands upon being heated, and contracts when temperatures are reduced. This can result in significant temperature gradients, especially during transient engine operation. The temperature gradients and differential thermal movement of the nozzle components result in thermally induced strain and stress that must be kept within suitable limits to ensure life expectancy of the nozzle.
The trailing edges of conventional vanes are particularly susceptible to thermal stress. Because it is very thin compared to the rest of the airfoil, the trailing edge responds more quickly to hot combustion gas flow than the surrounding airfoil material, resulting in severe temperature gradients that generate high thermal stress. Furthermore, the trailing edge is typically much hotter than the rest of the airfoil. Even with conventional cooling, the thermal stress can be sufficiently high to cause fatigue cracks in the trailing edge. Such cracks have an adverse impact on engine performance and may even result in engine failure should multiple cracks be allowed to link together.
Trailing edge distress can be reduced or eliminated by providing sufficient cooling to the vane trailing edge. Conventional cooling of a modem high pressure turbine nozzle vane is accomplished by film cooling from pressure side film cooling holes and pressure side slot film cooling. In addition, suction side film cooling holes also aid in cooling the trailing edge. However, aerodynamics on the airfoil are such that cooling air introduced on the suction side has a detrimental impact on turbine efficiency. In particular, introduction of film cooling air on the suction side just downstream of the nozzle throat plane is significantly detrimental to performance. Therefore, the majority of trailing edge cooling is provided by the pressure side film slots.
Thermal barrier coatings are commonly used to supplement existing impingement and/or film cooling means and thereby avoid trailing edge cracking. However, known thermal barrier coatings are relatively expensive and thus add to the cost of the nozzle vane.
Accordingly, there is a need for a turbine airfoil having improved trailing edge cooling that does not require a thermal barrier coating.
The above-mentioned need is met by the present invention which provides an airfoil having a pressure side and a suction side joined at a trailing edge wall that defines a trailing edge. The airfoil includes at least one cooling hole extending through the trailing edge wall so as to pass cooling fluid from the pressure side of the airfoil to the suction side.
The present invention and its advantages over the prior art will become apparent upon reading the following detailed description and the appended claims with reference to the accompanying drawings.